The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzle cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases which flow through turbine stages for extracting energy therefrom. In a turbofan engine, a high pressure turbine powers the compressor, and a low pressure turbine powers a fan disposed upstream from the compressor. Each turbine includes a stationary turbine nozzle having vanes mounted between inner and outer bands, followed in turn by a row of rotor blades extending outwardly from a rotor disk.
The high pressure turbine nozzle is disposed at the outlet of the combustor and receives therefrom combustion gases at the hottest temperature, with the temperature decreasing as energy is extracted from the gases in the downstream turbine stages. Both the nozzle vanes and rotor blades have hollow airfoils through which a portion of air bled from the compressor is used for providing cooling thereof. Bleeding cooling air from the compressor necessarily decreases the overall efficiency of the engine, and it is therefore desired to use as little cooling air as possible while adequately cooling the vanes and blades.
The vanes are stationary components whereas the blades are rotary components, and therefore have correspondingly different cooling arrangements in view of their different operating environment including centrifugal and thermal stress, and variations in heat transfer coefficient between the combustion gases and the respective airfoils thereof.
In one type of turbofan aircraft engine enjoying successful commercial use in this country for many years, the high pressure turbine nozzle includes three radial cooling channels between the leading and trailing edges of the vane airfoils, which are separated by corresponding radial ribs or bridges. Cooling air is provided in each vane through a common inlet in the outer band thereof, with a portion of the air splitting radially inwardly through a leading edge channel and an adjacent midchord channel. Except at the common outer inlet for the two channels, the cooling air is separately channeled therethrough.
The vanes include various film cooling holes through the pressure and suction sides thereof from which the air from the two channels is discharged for providing external film cooling of the vanes during operation.
The nozzle also includes a third, or trailing edge channel disposed aft of the midchord channel and separated therefrom by a corresponding rib or bridge. This third channel includes spaced apart pins between the pressure and suction sides of the vane for enhancing heat transfer of the cooling air channeled therebetween. Each vane includes a row of trailing edge outlet holes from which the cooling air in the third channel is discharged.
The third channel may receive its cooling air from its own inlet at the outer band, as well as an additional portion of the air from the second channel by truncating the midchord rib at its inner end. In an alternate embodiment, the midchord rib may include a row of crossover holes which provide the sole source of air into the third channel from the second channel along its span height.
To enhance the cooling effectiveness of the air channeled through the midchord channel, transversely extending turbulator ribs are typically disposed inside the pressure or concave side of the vane. Neither the first nor the third channels include turbulators to avoid the pressure drop associated therewith for maximizing engine performance while providing acceptable cooling.
The leading edges of the vanes typically have the most severe cooling requirements. They first receive the hot combustion gases which split along the pressure and suction sides of the vanes and effect significant stagnation pressure along the vane leading edges. The combustion gases have a high heat transfer coefficient along the vane leading edges and a high static pressure.
Accordingly, the air channeled through the leading edge cooling channel must have sufficient pressure greater than that of the external combustion gases to effect a backflow margin to prevent ingestion of the hot combustion gases through the film cooling holes and into the blades.
Pressure losses in the cooling air channeled through the vanes typically increase as the complexity of the cooling features increase. Although turbulators enhance cooling effectiveness they do so at the corresponding penalty and associated pressure losses therewith. This in turn requires that the provided cooling air have sufficient pressure for accommodating the expected losses therein for maintaining adequate backflow margin along the complete extent of the cooling channels to the last outlet hole.
Although the above described turbine nozzle has enjoyed many years of successful commercial use, a substantial power growth of the engine requires a corresponding increase in cooling of the nozzle which cannot be met by the present design. Engine power growth is being effected by a substantial increase in combustion gas temperature. The hotter combustion gases require a more effectively cooled turbine nozzle without excessively increasing the cooling air requirements from the compressor.
Accordingly, it is desired to provide a turbine nozzle having improved cooling features while minimizing pressure losses associated therewith.